Combustion turbines comprise a casing or cylinder for housing a compressor section, combustion section and turbine section. The compressor section comprises an inlet end and a discharge end. The combustion section or combustor comprises an inlet end and a combustor transition. The combustor transition is proximate the discharge end of the combustion section and comprises a wall which defines a flow channel which directs the working fluid into the turbine section's inlet end.
A supply of air is compressed in the compressor section and directed into the combustion section. Fuel enters the combustion section by means of a nozzle. The compressed air enters the combustion inlet and is mixed with the fuel. The air/fuel mixture is then combusted to produce high temperature and high pressure gas. This working gas is then ejected past the combustor transition and injected into the turbine section to run the turbine.
The turbine section comprises rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas flows through the turbine section causing the turbine blades to rotate, thereby turning the rotor, which is connected to a generator for producing electricity.
As those skilled in the art are aware, the maximum power output of a combustion turbine is achieved by heating the gas flowing through the combustion section to as high a temperature as is feasible. The hot gas, however, heats the various turbine components, such as the combustor, transition ducts, vanes and ring segments, that it passes when flowing through the turbine.
Accordingly, the ability to increase the combustion firing temperature is limited by the ability of the turbine components to withstand increased temperatures. Consequently, various cooling methods have been developed to cool turbine hot parts. These methods include open-loop air cooling techniques and closed-loop cooling systems. Both techniques, however, require significant design complexity, have considerable installation and operating costs and often carry attendant losses in turbine efficiency.
In addition, various insulation materials have been developed to strengthen the resistance of turbine critical components to increased temperature. Thermal Barrier Coatings (TBCs) are commonly used to protect critical components from premature breakdown due to increased temperatures to which the components are exposed. Generally, TBCs extend the life of critical components by reducing the rate of metal waste (through spalling) by oxidation.
In Advanced Turbine Systems (ATSs), however, the temperature demands of operation and the limits of ATS state-of-the-art materials, often lead to failure of the TBCs. This, in turn, results in premature failure of the critical components and therefore, failure of the turbine, interruption in the power supply and expensive repair costs. It is, therefore, desirable to provide turbine components that can withstand high temperatures without the use of thermal barrier coatings and reduce the need for cooling.
Commercially available ceramic matrix composites (CMCs) have many potential applications in gas turbines, but are limited in their exposure to temperatures near 1200.degree. C. for long periods of time, i.e., greater than 10,000 hours for gas turbines used in power generation. In addition, CMCs cannot be effectively cooled under high temperature conditions (&gt;1400.degree. C.) or high heat flux conditions due to their relatively low thermal conductivity and inability to fabricate intricate cooling passages.
Combustion of the fuel/air mixture occurs at temperatures much higher than the melting point of the metallic combustor liner. For this reason, the liners must be cooled by non-combusted, cooler air and are usually coated with thermal barrier coatings. The most common way of cooling metallic liners is by way of film cooling, which introduces cool air through the wall of the liner by way of small holes drilled at an acute angle to the surface. This air, in turn, forms a cooler boundary layer on the inside surface of the combustor liner, protecting it from the hot combustion gases. One of the problems with film cooling is that undesirable combustion byproducts (carbon monoxide (CO) and unburned hydrocarbons (UHC)) occur when the cooler air mixes with the hot gases. In anticipation of dilution due to film cooling, the fuel/air mixture is consequently richer than desirable, resulting in excessive NO.sub.x emissions. A true hot wall combustor requires no film cooling (resulting in lower CO and UHC emissions), allows leaner combustion (resulting in lower NO.sub.x emissions), and provides increased flame stability (resulting in greater durability and reliability).
The transition duct is a large, complex structure which contains the hot combustion gases and directs them into the turbine inlet. The large surface area and the high internal temperature make these parts extremely difficult to cool effectively. Conventional transitions are made from Nickel-based superalloys coated internally with thermal barrier coatings. The latest high efficiency utility engines necessitate that these parts be actively cooled, requiring internal wall cooling passages, and complex and costly construction. With much simpler construction, lower cost components would be possible using an insulated CMC concept. Passive cooling methods could be employed using redirected combustor inlet gases, resulting in net efficiency gains.
The first stage of turbine vanes direct the combustion exhaust gases to the airfoil portions of the first row of rotating turbine blades. These vanes are subjected to high velocity, high temperature gases under high pressure conditions. In addition, these are complex parts with high surface areas and, therefore, are difficult to cool to acceptable temperatures. Conventional state-of-the-art first row turbine vanes are fabricated from single-crystal superalloy castings with intricate cooling passages and with external thermal barrier coatings applied. Not only are these components expensive to manufacture, but with ever-increasing gas path temperatures, their ability to be effectively cooled is limited. Higher temperature materials would obviate the need for such complexity, thus minimizing cost, and also minimizing the need for cooling air, thereby improving engine efficiency and reducing operating costs.
The rotating turbine or rotor of an axial flow gas turbine consists of a plurality of blades attached to a rotor disk. In operation, the shaft and blades rotate inside a shroud. Preferably, the inner surface of the inner wall of the shroud is coated with an abradable material. The initial placement of the rotor blades are such that the blade tips are as close as possible to the coating.
Materials which abrade readily in a controlled fashion are used in a variety of applications. One such material is disclosed in European Patent Office publication No. 007,511,04, entitled "An Abradable Composition," filed Jan. 2, 1997, which is incorporated herein by reference in its entirety. Contact between a rotating part and a fixed abradable seal causes the abradable material to wear in a configuration which closely mates with and conforms to the moving part at the region of contact. The moving part wears away a portion of the abradable seal so that the seal assumes a geometry which precisely fits the moving part, i.e., a close clearance gap. This effectively forms a seal having an extremely close tolerance.
As appreciated by those skilled in the art, it is important to reduce leakage in axial flow gas turbines to maximize turbine efficiency. This is achieved by minimizing the clearance between the blade tips and the inner wall of the shroud. As the turbine blades rotate, however, they expand slightly due to the heat generated by the turbine. The tips of the rotating blades then contact the abradable material and carve precisely-defined grooves in the coating without contacting the shroud itself. These grooves provide for the blades to rotate, resulting in a customized seal for the turbine blades. It is, therefore, desirable to provide an abradable material that abrades relatively easily without wearing down the blade tips.
Abradable materials are also used for high temperature insulation. Abradability is usually achieved by altering the density of the material by introducing microscopic porosity. The consequence of this, however, is a reduction in the erosion resistance of the abradable coating. Alternatively, coatings can be fabricated with higher densities for acceptable erosion resistance. This, in turn, sacrifices abradability, necessitating the use abrasive blade tip treatments. Relatively low thermal conductivity and relatively high erosion resistance are two properties of abradable materials required for high temperature insulation. These characteristics are especially important in an ATS environment, where temperatures can approach 1600.degree. C. It is, therefore, desirable to provide an abradable material that has relatively low thermal conductivity and relatively high erosion resistance, particularly at elevated temperatures.
European Patent Office publication No. 007,511,04 discloses a ceramic abradable material that can be used to seal ceramic turbine components. This material, however, purportedly has a high temperature capability of only 1300.degree. C., not suitable for use in ATS turbines. It is, therefore, desirable to provide a ceramic abradable material that can be used in ATS turbines, where temperatures can approach 1600.degree. C.